The present invention relates to materials designed to withstand high temperatures. More particularly, this invention relates to heat-resistant alloys for high-temperature applications, such as, for instance, gas turbine engine components of aircraft engines and power generation equipment.
There is a continuing demand in many industries, notably in the aircraft engine and power generation industries where efficiency directly relates to operating temperature, for alloys that exhibit sufficient levels of strength and oxidation resistance at increasingly higher temperatures. Gas turbine airfoils on such components as vanes and blades are usually made of materials known in the art as xe2x80x9csuperalloys.xe2x80x9d The term xe2x80x9csuperalloyxe2x80x9d is usually intended to embrace iron-, cobalt-, or nickel-based alloys, which include one or more additional elements to enhance high temperature performance, including such non-limiting examples as aluminum, tungsten, molybdenum, titanium, and iron. The term xe2x80x9cbasedxe2x80x9d as used in, for example, xe2x80x9cnickel-based superalloyxe2x80x9d is widely accepted in the art to mean that the element upon which the alloy is xe2x80x9cbasedxe2x80x9d is the single largest elemental component by weight in the alloy composition. Generally recognized to have service capabilities limited to a temperature of about 1100xc2x0 C., conventional superalloys used in gas turbine airfoils often operate at the upper limits of their practical service temperature range. In jet engines, for example, bulk average airfoil temperatures range from about 900xc2x0 C. to about 1000xc2x0 C., while airfoil leading and trailing edge and tip temperatures can reach about 1150xc2x0 C. or more. At such elevated temperatures, the oxidation process consumes conventional superalloy parts, forming a weak, brittle metal oxide that is prone to chip or spall away from the part.
Erosion and oxidation of material at the edges of airfoils lead to degradation of turbine efficiency. As airfoils are worn away, gaps between components become excessively wide, allowing gas to leak through the turbine stages without the flow of the gas being converted into mechanical energy. When efficiency drops below specified levels, the turbine must be removed from service for overhaul and refurbishment. A significant portion of this refurbishment process is directed at the repair of the airfoil leading and trailing edges and tips. For example, damaged material is removed and then new material built onto the blade by welding with filler material or by laser deposition of metal powders. The performance of alloys commonly used for repair is comparable or inferior to that of the material of the original component, depending upon the microstructure of the repaired material, its defect density due to processing, and its chemistry. Furthermore, in current practice, the original edge material is made of the same material as the rest of the original blade, often a superalloy based on nickel or cobalt. Because this material was selected to balance the design requirements of the entire blade, it is generally not optimized to meet the special local requirements demanded by conditions at the airfoil leading or trailing edges. However, maximum temperatures, such as those present at airfoil tips and edges, are expected in future applications to be over about 1300xc2x0 C., at which point many conventional superalloys begin to melt. Clearly, new materials for repair and manufacture must be developed to improve the performance of repaired components and to exploit efficiency enhancements available to new components designed to operate at higher turbine operating temperatures.
Maximum temperatures are expected in future applications to be over about 1300xc2x0 C., at which point many conventional superalloys begin to melt. Clearly, new materials must be developed if the efficiency enhancements available at higher operating temperatures are to be exploited.
The so-called xe2x80x9crefractory superalloys,xe2x80x9d as described in Koizumi et al., U.S. Pat. No. 6,071,470, represent a class of alloys designed to operate at higher temperatures than those of conventional superalloys. According to Koizumi et al., refractory superalloys consist essentially of a primary constituent selected from the group consisting of iridium (Ir), rhodium (Rh), and a mixture thereof, and one or more additive elements selected from the group consisting of niobium (Nb), tantalum (Ta), hafnium (Hf), zirconium (Zr), uranium (U), vanadium (V), titanium (Ti), and aluminum (Al). The refractory superalloys have a microstructure containing an FCC (face-centered cubic)-type crystalline structure phase and an L12 type crystalline structure phase, and the one or more additive elements are present in a total amount within the range of from 2 atom % to 22 atom %.
Although the refractory superalloys have shown potential to become replacements for conventional superalloys in present and future gas turbine engine designs, it has been shown that many alloys of this class do not meet all of the desired performance criteria for high-temperature applications. Therefore, the need persists for alloys with improved high-temperature properties.
The present invention provides several embodiments that address this need. One embodiment provides an alloy comprising ruthenium, nickel, aluminum, and chromium, wherein a microstructure of the alloy is essentially free of an L12-structured phase at temperatures greater than about 1000xc2x0 C. and comprises an A3-structured phase and up to about 40 volume percent of a B2-structured phase.
A second embodiment provides a gas turbine engine component comprising: an alloy comprising ruthenium, nickel, aluminum, and chromium, wherein a microstructure of the alloy of the engine component is essentially free of an L12-structured phase at temperatures greater than about 1000xc2x0 C. and comprises an A3-structured phase and up to about 40 volume percent of a B2-structured phase.
A third embodiment provides a repair material for repairing superalloy articles, the repair material comprising ruthenium, nickel, aluminum, and chromium, wherein a microstructure of the repair material is essentially free of an L12-structured phase at temperatures greater than about 1000xc2x0 C. and comprises an A3-structured phase and up to about 40 volume percent of a B2-structured phase.
A fourth embodiment provides a repaired gas turbine engine component comprising a repair material comprising ruthenium, nickel, aluminum, and chromium, wherein a microstructure of the repair material of the engine component is essentially free of an L12-structured phase at temperatures greater than about 1000xc2x0 C. and comprises an A3-structured phase and up to about 40 volume percent of a B2-structured phase.
A fifth embodiment provides a method for repairing a gas turbine engine component, the method comprising providing a repair material, the repair material comprising ruthenium, nickel, aluminum, and chromium, wherein a microstructure of the repair material is essentially free of an L12-structured phase at temperatures greater than about 1000xc2x0 C. and comprises an A3-structured phase and up to about 40 volume percent of a B2-structured phase; and joining the repair material to the gas turbine engine component.